Field of the Invention
The field of the invention relates to turbines generally, and more particularly to certain new and useful advances in the manufacture and/or cooling of gas turbine combustor liners, of which the following is a specification, reference being had to the drawings accompanying and forming a part of the same.
Description of Related Art
A combustor of a gas turbine is a component or area thereof in which combustion of fuel occurs, and which affects various engine characteristics, including emissions and/or fuel efficiency. The purpose of combustors is to regulate the combustion of fuel and air to produce energy in the form of high-temperature gases, which can rotate an engine or generator turbine and/or be routed through an exhaust nozzle. Combustors are subject to various design considerations, which include, but are not limited to: maintaining a uniform exit temperature profile so that hot spots do not damage the turbine or the combustor, and operating with low emission of pollutants. Accordingly, a combustor liner, which contains the combustion process and introduces various airflows into the combustion zone, is built to withstand high temperatures. Some combustor liners are insulated from heat by thermal barrier coatings (“TBCs”), but most rely on various types of air-cooling to reduce liner temperature. For example, film cooling injects a thin blanket of cool air over the interior of the combustor liner, while effusion cooling pushes cool air through a lattice formed of closely spaced, discrete pores, or holes, in the combustor liner. Of the two approaches, effusion cooling tends to use less air and to generate a more uniform temperature profile than film cooling.
FIG. 14 is a sectional side view of a substrate coated with a thermal barrier coating and having a conventional round cooling hole 120. FIG. 15 is another sectional view of the conventional round cooling hole 120 of FIG. 14, taken along the line A-A′. FIG. 16 is another sectional view of the conventional round cooling hole 120 of FIG. 15, taken along the line B-B′.
FIG. 17 is a sectional side view of a substrate coated with a thermal barrier coating and having a conventional conical film cooling hole 130. FIG. 18 is another sectional view of the conventional conical film cooling hole 130 of FIG. 17, taken along the line A-A′. FIG. 19 is another sectional view of the conventional conical film cooling hole 130 of FIG. 17, taken along the line B-B′.
FIG. 20 is a sectional side view of a substrate coated with a thermal barrier coating and having a conventional “3D” film cooling hole 140. FIG. 21 is another sectional view of the conventional “3D” film cooling hole 140 of FIG. 20, taken along the line A-A′. FIG. 22 is another sectional view of the conventional “3D” film cooling hole 140 of FIG. 20, taken along the line B-B′.
FIG. 23 is a sectional side view of a substrate coated with a thermal barrier coating and having a conventional “fan” film cooling hole 150. FIG. 24 is another sectional view of the conventional “fan” film cooling hole 150 of FIG. 23, taken along the line A-A′. FIG. 25 is another sectional view of the conventional “fan” film cooling hole 150 of FIG. 23, taken along the line B-B′.
Referring to FIGS. 15-25, each conventional cooling hole 120, 130, 140 and 150 is formed at an angle in a substrate 100. The substrate 100 is coated with a thermal barrier coating 101. The thermal barrier coating 101 is coated with a bond coat 103. Each cooling hole 120, 130, 140 and 150 has an inlet 113 formed on one side of the substrate 100 and a larger outlet 111 that is formed on the opposite side of the substrate 100. Each cooling hole 120 130, 140 and 150 has a bore 112 that communicates with and/or forms part of the inlet 113. The bore 112 is generally cylindrical. For the round cooling hole 120, the diameter 114 of the bore 112 is uniform between the inlet 113 and the outlet 112. For the cooling holes 130, 140 and 150, the diameter 114 of the bore 112 increases proximate the outlet 111.
However, each of the convention cooling holes 120, 130, 140 and 150 has at least one disadvantage. For example, analyses of the conical film cooling holes 130 and of the “fan” film cooling holes 150 has revealed drawbacks in convective cooling. As shown, the “3D” film cooling holes 140 have cylindrical bores 112 that transition to three-dimensional diffusion on all sides in the downstream direction. However, this type of effusion cooling arrangement tends to be unsuitable for combustor liners because such three-dimensional downstream diffusion removes a significant amount of thermal barrier coating (“TBC”) from the combustor liner, a disadvantage in combustors where radiation is a substantial part of the heat load.
The practice in effusion cooling has been to limit the axial and radial spacing of multihole arrays to about 6.5 diameters to ensure the respective airflows coalesce into a continuous protective film and to ensure every location has bore convective cooling. This spacing implies a certain minimum cooling flow per unit area. However, as technology advances, there is a strong desire to reduce the cooling flow and free up air for reduced NOx emissions, increased efficiency, and/or better turbine cooling.
Similarly there has been a limit on the axial and tangential spacing of multihole arrays of 4 or 5 diameters minimum to avoid excessive stress concentrations. This spacing implies a certain maximum cooling flow per unit area. At certain locations, however, where the accumulation of film is disrupted by local flow characteristic within the combustor, there is a strong desire to locally increase the cooling flow to avoid low life of the liner. Enlarging the inlet metering hole in the direction of the sides gives the best supply of air to the shape wings and gives more bore cooling surface area than simply enlarging the metering hole diameter.